This invention relates to gas turbine engines, and, more particularly, to the structure and fabrication of turbine blades used in such engines.
In a gas turbine engine such as used by jet aircraft, air is drawn into the front of the engine and compressed by a shaft-mounted compressor. The compressed air is mixed with fuel, and the mixture is burned in a combustor. The resulting hot exhaust gases are passed through a turbine that causes the compressor shaft to turn, and then out the rear of the engine to provide forward thrust.
The turbine of the gas turbine engine includes stationary turbine vanes that redirect the generally axial flow of hot exhaust gas so that it has a small sideways momentum component. Turbine blades are mounted on a turbine disk that in turn is mounted on a rotating shaft, usually the same shaft that turns the compressor. The impact of the hot exhaust gas on the turbine blades forces the turbine blades to rotate circumferentially. The turbine disk turns, driving the compressor through the rotation of the shaft.
The turbine blades and turbine vanes are key components of the gas turbine engine. They must operate in a high-temperature oxidizing environment, which may carry salt and other corrosive and erosive agents. The turbine blades operate under high stresses created by centrifugal forces as the turbine disk turns. The ability of the turbine blades and turbine vanes to operate in these conditions is essential to the efficient operation of the gas turbine engine, because engine efficiency increases with increasing temperature of the hot exhaust gas that enters the turbine section of the engine. The turbine blades of the high pressure section and the forward stages of the low pressure section of the turbine section experience the highest temperatures and operating stresses. In most cases, the performance of these turbine blades limits the performance of the engine. The present invention is of most direct benefit to these turbine blades, but is applicable to the other turbine blades in the turbine section as well.
Because of the critical importance of the turbine blades to the operation of the gas turbine engine, a great deal of attention has been directed to their improvement. Turbine blades are typically made of complex alloys of nickel, cobalt, and other elements. Alloys of this type have been developed specifically for this application. The turbine blades may be made as single crystals or directionally aligned polycrystals, to take advantage of the mechanical properties of particular crystallographic directions. The turbine blades are often coated with thermal barrier coating systems to increase the permissible temperature of the hot exhaust gas. The turbine blades also may have cooling channels therethrough so that cooling air can be passed through the interior of the blades to reduce the blade temperature.
Although turbine blades must be engineered to operate in these highly adverse environments, they must also be manufactured in a reasonably economical manner. A typical commercial gas turbine engine may contain 80 turbine blades or more, and the cost of the turbine blades can add a substantial amount to the initial cost of the engine and its repair costs as the turbine blades are replaced during the operating life of the engine. Thus, there is a continuing need for improved, reduced cost manufacturing techniques to fabricate complex gas turbine blades from superalloy materials and with the required crystallographic structures and orientations.